Turbofan engine having angled inlet pre-swirl vanes

ABSTRACT

A turbofan engine is provided. The turbofan engine includes a fan having a plurality of fan blades; a turbomachine operably coupled to the fan for driving the fan, the turbomachine having a compressor section, a combustion section, and a turbine section in serial flow order and together defining a core air flowpath; a nacelle surrounding and at least partially enclosing the fan, the nacelle defining a radius and a longitudinal axis; and an inlet pre-swirl vane located upstream of the plurality of fan blades and defining a chord, the inlet pre-swirl vane coupled to the nacelle, wherein the inlet pre-swirl vane is angled at a first angle with respect to the radius of the nacelle, and wherein the chord of the inlet pre-swirl vane is angled at a second angle with respect to the longitudinal axis of the nacelle.

TECHNICAL FIELD

The present subject matter relates generally to a gas turbine engine, ormore particularly to a gas turbine engine configured to guide an airflowat an inlet of a nacelle and to direct incoming objects towards an outerportion of the engine and away from a core of the engine.

BACKGROUND

A turbofan engine generally includes a fan having a plurality of fanblades and a turbomachine arranged in flow communication with oneanother. Additionally, the turbomachine of the turbofan engine generallyincludes, in serial flow order, a compressor section, a combustionsection, a turbine section, and an exhaust section. In operation, air isprovided from the fan to an inlet of the compressor section where one ormore axial compressors progressively compress the air until it reachesthe combustion section. Fuel is mixed with the compressed air and burnedwithin the combustion section to provide combustion gases. Thecombustion gases are routed from the combustion section to the turbinesection. The flow of combustion gasses through the turbine sectiondrives the turbine section and is then routed through the exhaustsection, e.g., to atmosphere.

BRIEF DESCRIPTION OF THE DRAWINGS

A full and enabling disclosure of the present disclosure, including thebest mode thereof, directed to one of ordinary skill in the art, is setforth in the specification, which makes reference to the appendedfigures, in which:

FIG. 1 is a schematic cross-sectional view of an exemplary gas turbineengine according to an exemplary embodiment of the present subjectmatter.

FIG. 2 is a close-up, schematic, cross-sectional view of a forward endof the exemplary gas turbine engine of FIG. 1 according to an exemplaryembodiment of the present subject matter.

FIG. 3 is a schematic view of an inlet to the exemplary gas turbineengine of FIG. 1 , along an axial direction of the gas turbine engine ofFIG. 1 according to an exemplary embodiment of the present subjectmatter.

FIG. 4 it is a schematic view of an inlet to a gas turbine engine inaccordance with another exemplary embodiment of the present disclosure.

FIG. 5 is a schematic view of an inlet to the exemplary gas turbineengine of FIG. 1 , along an axial direction of the gas turbine engine ofFIG. 1 according to another exemplary embodiment of the present subjectmatter.

FIG. 6 is a cross-sectional view of a part span inlet guide vane of theexemplary gas turbine engine of FIG. 1 at a first location along a spanof the part span inlet guide vane.

FIG. 7 is a cross-sectional view of the part span inlet guide vane ofthe exemplary gas turbine engine of FIG. 1 at a second location alongthe span of the part span inlet guide vane.

FIG. 8 is a schematic view of an inlet to the exemplary gas turbineengine of FIG. 1 , along an axial direction of the gas turbine engine ofFIG. 1 according to another exemplary embodiment of the present subjectmatter.

FIG. 9 is a schematic view of an inlet to the exemplary gas turbineengine of FIG. 1 , along an axial direction of the gas turbine engine ofFIG. 1 according to another exemplary embodiment of the present subjectmatter.

FIG. 10 is a schematic view of an inlet to the exemplary gas turbineengine of FIG. 1 , along an axial direction of the gas turbine engine ofFIG. 1 according to another exemplary embodiment of the present subjectmatter.

FIG. 11 is a schematic view of an inlet to the exemplary gas turbineengine of FIG. 1 , along an axial direction of the gas turbine engine ofFIG. 1 according to another exemplary embodiment of the present subjectmatter.

Corresponding reference characters indicate corresponding partsthroughout the several views. The exemplifications set out hereinillustrate exemplary embodiments of the disclosure, and suchexemplifications are not to be construed as limiting the scope of thedisclosure in any manner.

DETAILED DESCRIPTION

Reference will now be made in detail to present embodiments of thedisclosure, one or more examples of which are illustrated in theaccompanying drawings. The detailed description uses numerical andletter designations to refer to features in the drawings. Like orsimilar designations in the drawings and description have been used torefer to like or similar parts of the disclosure.

The following description is provided to enable those skilled in the artto make and use the described embodiments contemplated for carrying outthe disclosure. Various modifications, equivalents, variations, andalternatives, however, will remain readily apparent to those skilled inthe art. Any and all such modifications, variations, equivalents, andalternatives are intended to fall within the scope of the presentdisclosure.

The word “exemplary” is used herein to mean “serving as an example,instance, or illustration.” Any implementation described herein as“exemplary” is not necessarily to be construed as preferred oradvantageous over other implementations. Additionally, unlessspecifically identified otherwise, all embodiments described hereinshould be considered exemplary.

For purposes of the description hereinafter, the terms “upper”, “lower”,“right”, “left”, “vertical”, “horizontal”, “top”, “bottom”, “lateral”,“longitudinal”, and derivatives thereof shall relate to the disclosureas it is oriented in the drawing figures. However, it is to beunderstood that the disclosure may assume various alternativevariations, except where expressly specified to the contrary. It is alsoto be understood that the specific devices illustrated in the attacheddrawings, and described in the following specification, are simplyexemplary embodiments of the disclosure. Hence, specific dimensions andother physical characteristics related to the embodiments disclosedherein are not to be considered as limiting.

As used herein, the terms “first”, “second”, and “third” may be usedinterchangeably to distinguish one component from another and are notintended to signify location or importance of the individual components.

The terms “forward” and “aft” refer to relative positions within a gasturbine engine, with forward referring to a position closer to an engineinlet and aft referring to a position closer to an engine nozzle orexhaust.

The terms “upstream” and “downstream” refer to the relative directionwith respect to fluid flow in a fluid pathway. For example, “upstream”refers to the direction from which the fluid flows, and “downstream”refers to the direction to which the fluid flows.

The singular forms “a”, “an”, and “the” include plural references unlessthe context clearly dictates otherwise.

Additionally, the terms “low,” “high,” or their respective comparativedegrees (e.g., lower, higher, where applicable) each refer to relativespeeds or pressures within an engine, unless otherwise specified. Forexample, a “low-pressure turbine” operates at a pressure generally lowerthan a “high-pressure turbine.” Alternatively, unless otherwisespecified, the aforementioned terms may be understood in theirsuperlative degree. For example, a “low-pressure turbine” may refer tothe lowest maximum pressure turbine within a turbine section, and a“high-pressure turbine” may refer to the highest maximum pressureturbine within the turbine section. An engine of the present disclosuremay also include an intermediate pressure turbine, e.g., an enginehaving three spools.

Approximating language, as used herein throughout the specification andclaims, is applied to modify any quantitative representation that couldpermissibly vary without resulting in a change in the basic function towhich it is related. Accordingly, a value modified by a term or terms,such as “about”, “approximately”, and “substantially”, are not to belimited to the precise value specified. In at least some instances, theapproximating language may correspond to the precision of an instrumentfor measuring the value, or the precision of the methods or machines forconstructing or manufacturing the components and/or systems. Forexample, the approximating language may refer to being within a 1, 2, 4,10, 15, or 20 percent margin. These approximating margins may apply to asingle value, either or both endpoints defining numerical ranges, and/orthe margin for ranges between endpoints.

Here and throughout the specification and claims, range limitations arecombined and interchanged, such ranges are identified and include allthe sub-ranges contained therein unless context or language indicatesotherwise. For example, all ranges disclosed herein are inclusive of theendpoints, and the endpoints are independently combinable with eachother.

As used herein, the term “fan pressure ratio” refers to a ratio of anair pressure immediately downstream of the fan blades if a fan duringoperation of the fan to an air pressure immediately upstream of the fanblades of the fan during operation of the fan.

As used herein, the term “rated speed” with reference to a turbofanengine refers to a maximum rotational speed that the turbofan engine mayachieve while operating properly. For example, the turbofan engine maybe operating at the rated speed during maximum load operations, such asduring takeoff operations.

Also as used herein, the term “fan tip speed” as defined by theplurality of fan blades of the fan refers to a linear speed of an outertip of a fan blade along a radial direction during operation of the fan.

The present disclosure is generally related to an inlet pre-swirl vaneconfigured as a plurality of part span inlet guide vanes for a turbofanengine. In the present disclosure, the plurality of part span inletguide vanes are angled at a first angle with respect to a radius of anouter nacelle of the turbofan engine. Furthermore, the plurality of partspan inlet guide vanes define a chord and the chord of the part spaninlet guide vane is angled at a second angle with respect to thelongitudinal axis of the outer nacelle.

In this manner, the plurality of part span inlet guide vanes areconfigured to direct incoming objects towards an outer portion of theturbofan engine. The plurality of part span inlet guide vanes areconfigured to direct incoming objects away from a core air flowpath ofthe turbofan engine and towards a bypass airflow passage. This providesa deflection mechanism that facilitates ingestion of an object into anouter portion of the turbofan engine by minimizing the chance that theobject travels to the core of the turbofan engine. Such objects mayinclude bird strikes, hail, ice, sandstorms, debris, and other foreignobjects.

Furthermore, in this manner, the plurality of part span inlet guidevanes are also configured to pre-swirl an airflow provided through aninlet of the outer nacelle, upstream of the plurality of fan blades of afan. As discussed herein, pre-swirling the airflow provided through theinlet of the outer nacelle prior to such airflow reaching the pluralityof fan blades of the fan may reduce separation losses and/or shocklosses, allowing the fan to operate with relatively high fan tip speedswith less losses in efficiency. Furthermore, in this manner, theplurality of part span inlet guide vanes are also configured to minimizeflutter and maintain laminar air flow over the part span inlet guidevanes and minimize turbulence in the inlet.

Referring now to the drawings, wherein identical numerals indicate thesame elements throughout the figures, FIG. 1 is a schematiccross-sectional view of a gas turbine engine in accordance with anexemplary embodiment of the present disclosure. More particularly, forthe embodiment of FIG. 1 , the gas turbine engine is an aeronautical,turbofan jet engine 10, referred to herein as “turbofan engine 10”,configured to be mounted to an aircraft, such as in an under-wingconfiguration or tail-mounted configuration. As shown in FIG. 1 , theturbofan engine 10 defines an axial direction A (extending parallel to alongitudinal centerline 12 provided for reference), a radial directionR, and a circumferential direction (i.e., a direction extending aboutthe axial direction A; not depicted). In general, the turbofan engine 10includes a fan section 14 and a turbomachine 16 disposed downstream fromthe fan section 14 (the turbomachine 16 sometimes also, oralternatively, referred to as a “core turbine engine”).

The exemplary turbomachine 16 depicted generally includes asubstantially tubular outer casing 18 that defines an annular inlet 20.The outer casing 18 encases, in serial flow relationship, a compressorsection including a first, booster or low pressure (LP) compressor 22and a second, high pressure (HP) compressor 24; a combustion section 26;a turbine section including a first, high pressure (HP) turbine 28 and asecond, low pressure (LP) turbine 30; and a jet exhaust nozzle section32. A high pressure (HP) shaft 34 drivingly connects the HP turbine 28to the HP compressor 24. A low pressure (LP) shaft 36 drivingly connectsthe LP turbine 30 to the LP compressor 22. The compressor section,combustion section 26, turbine section, and jet exhaust nozzle section32 are arranged in serial flow order and together define a core airflowpath 37 through the turbomachine 16. It is also contemplated thatthe present disclosure is compatible with an engine having anintermediate pressure turbine, e.g., an engine having three spools.

Referring still to the embodiment of FIG. 1 , the fan section 14includes a variable pitch, single stage fan 38, the turbomachine 16operably coupled to the fan 38 for driving the fan 38. The fan 38includes a plurality of rotatable fan blades 40 coupled to a disk 42 ina spaced apart manner. As depicted, the fan blades 40 extend outwardlyfrom disk 42 generally along the radial direction R. Each fan blade 40is rotatable relative to the disk 42 about a pitch axis P by virtue ofthe fan blades 40 being operatively coupled to a suitable actuationmember 44 configured to collectively vary the pitch of the fan blades40, e.g., in unison. The fan blades 40, disk 42, and actuation member 44are together rotatable about the longitudinal centerline 12 by LP shaft36 across a power gear box 46. The power gear box 46 includes aplurality of gears for stepping down the rotational speed of the LPshaft 36 to a more efficient rotational fan speed. Accordingly, for theembodiment depicted, the turbomachine 16 is operably coupled to the fan38 through the power gear box 46.

In exemplary embodiments, the fan section 14 includes twenty-two (22) orfewer fan blades 40. In certain exemplary embodiments, the fan section14 includes twenty (20) or fewer fan blades 40. In certain exemplaryembodiments, the fan section 14 includes eighteen (18) or fewer fanblades 40. In certain exemplary embodiments, the fan section 14 includessixteen (16) or fewer fan blades 40. In certain exemplary embodiments,it is contemplated that the fan section 14 includes other number of fanblades 40 for a particular application.

During operation of the turbofan engine 10, the fan 38 defines a fanpressure ratio and the plurality of fan blades 40 each define a fan tipspeed. The exemplary turbofan engine 10 depicted defines a relativelyhigh fan tip speed and relatively low fan pressure ratio duringoperation of the turbofan engine at a rated speed. As used herein, theterm “fan pressure ratio” refers to a ratio of an air pressureimmediately downstream of the fan blades 40 during operation of the fan38 to an air pressure immediately upstream of the fan blades 40 duringoperation of the fan 38. For the embodiment depicted in FIG. 1 , the fan38 of the turbofan engine 10 defines a relatively low fan pressureratio. For example, the turbofan engine 10 depicted defines a fanpressure ratio less than or equal to about 1.5. For example, in certainexemplary embodiments, the turbofan engine 10 may define a fan pressureratio less than or equal to about 1.4. In certain exemplary embodiments,it is contemplated that the turbofan engine 10 may define other fanpressure ratios for a particular application. The fan pressure ratio maybe the fan pressure ratio of the fan 38 during operation of the turbofanengine 10, such as during operation of the turbofan engine 10 at a ratedspeed.

As used herein, the term “rated speed” with reference to the turbofanengine 10 refers to a maximum rotational speed that the turbofan engine10 may achieve while operating properly. For example, the turbofanengine 10 may be operating at the rated speed during maximum loadoperations, such as during takeoff operations.

Also as used herein, the term “fan tip speed” defined by the pluralityof fan blades 40 refers to a linear speed of an outer tip of a fan blade40 along the circumferential direction during operation of the fan 38.In exemplary embodiments, the turbofan engine 10 of the presentdisclosure causes the fan blades 40 of the fan 38 to rotate at arelatively high rotational speed. For example, during operation of theturbofan engine 10 at the rated speed, the fan tip speed of each of theplurality of fan blades 40 is greater than or equal to 1,000 feet persecond and less than or equal to 2,250 feet per second. In certainexemplary embodiments, during operation of the turbofan engine 10 at therated speed, the fan tip speed of each of the fan blades 40 may begreater than or equal to 1,250 feet per second and less than or equal to2,250 feet per second. In certain exemplary embodiments, duringoperation of the turbofan engine 10 at the rated speed, the fan tipspeed of each of the fan blades 40 may be greater than or equal to about1,350 feet per second, such as greater than about 1,450 feet per second,such as greater than about 1,550 feet per second, and less than or equalto 2,250 feet per second. In certain exemplary embodiments, it iscontemplated that during operation of the turbofan engine 10 at therated speed, the fan tip speed of each of the fan blades 40 may defineother ranges for a particular application.

Referring still to the exemplary embodiment of FIG. 1 , the disk 42 iscovered by rotatable front nacelle or hub 48 aerodynamically contouredto promote an airflow through the plurality of fan blades 40.Additionally, the exemplary fan section 14 includes an annular fancasing or outer nacelle 50 that at least partially, and for theembodiment depicted, circumferentially, surrounds the fan 38 and atleast a portion of the turbomachine 16.

More specifically, the outer nacelle 50 includes an inner wall 52 and adownstream section 54 of the inner wall 52 of the outer nacelle 50extends over an outer portion of the turbomachine 16 so as to define abypass airflow passage 56 therebetween. Additionally, for the embodimentdepicted, the outer nacelle 50 is supported relative to the turbomachine16 by a plurality of circumferentially spaced outlet guide vanes 55. Theouter nacelle 50 includes an inlet 60 at a leading edge 61 of the outernacelle 50.

During operation of the turbofan engine 10, a volume of air 58 entersthe turbofan engine 10 through the inlet 60 of the outer nacelle 50and/or fan section 14. As the volume of air 58 passes across the fanblades 40, a first portion of the air 58 as indicated by arrows 62 isdirected or routed into the bypass airflow passage 56 and a secondportion of the air 58 as indicated by arrow 64 is directed or routedinto the core air flowpath 37. The ratio between an amount of airflowthrough the bypass airflow passage 56 (i.e., the first portion of airindicated by arrows 62) to an amount of airflow through the core airflowpath 37 (i.e., the second portion of air indicated by arrows 64) isknown as a bypass ratio.

Referring still to FIG. 1 , the compressed second portion of airindicated by arrows 64 from the compressor section mixes with fuel andis burned within the combustion section to provide combustion gases 66.The combustion gases 66 are routed from the combustion section 26,through the HP turbine 28 where a portion of thermal and/or kineticenergy from the combustion gases 66 is extracted via sequential stagesof HP turbine stator vanes 68 that are coupled to the outer casing 18and HP turbine rotor blades 70 that are coupled to the HP shaft 34, thuscausing the HP shaft 34 to rotate, thereby supporting operation of theHP compressor 24. The combustion gases 66 are then routed through the LPturbine 30 where a second portion of thermal and kinetic energy isextracted from the combustion gases 66 via sequential stages of LPturbine stator vanes 72 that are coupled to the outer casing 18 and LPturbine rotor blades 74 that are coupled to the LP shaft 36, thuscausing the LP shaft 36 to rotate, thereby supporting operation of theLP compressor 22 and/or rotation of the fan 38.

The combustion gases 66 are subsequently routed through the jet exhaustnozzle section 32 of the turbomachine 16 to provide propulsive thrust.Simultaneously, the pressure of the first portion of air indicated byarrows 62 is substantially increased as the first portion of air 62 isrouted through the bypass airflow passage 56 before it is exhausted froma fan nozzle exhaust section 76 of the turbofan 10, also providingpropulsive thrust. The HP turbine 28, the LP turbine 30, and the jetexhaust nozzle section 32 at least partially define a hot gas path 78for routing the combustion gases 66 through the turbomachine 16.

Referring still to FIG. 1 , the turbofan engine 10 of the presentdisclosure also provides pre-swirling flow forward of a tip of the fanblade 40 as described herein. For example, the turbofan engine 10additionally includes an inlet pre-swirl vane, e.g., configured as aplurality of part span inlet guide vanes 100, as described in greaterdetail herein.

In some exemplary embodiments, it will be appreciated that the exemplaryturbofan engine 10 of the present disclosure may be a relatively largepower class turbofan engine 10. Accordingly, when operated at the ratedspeed, the turbofan engine 10 may be configured to generate a relativelylarge amount of thrust. More specifically, when operated at the ratedspeed, the turbofan engine 10 may be configured to generate at leastabout 20,000 pounds of thrust, such as at least about 25,000 pounds ofthrust, such as at least about 30,000 pounds of thrust, and up to, e.g.,about 150,000 pounds of thrust. Accordingly, the turbofan engine 10 maybe referred to as a relatively large power class gas turbine engine.

Moreover, it should be appreciated that the exemplary turbofan engine 10depicted in FIG. 1 is by way of example only, and that in otherexemplary embodiments, the turbofan engine 10 may have any othersuitable configuration. For example, in certain exemplary embodiments,the fan may not be a variable pitch fan, the engine may not include areduction gearbox (e.g., power gearbox 46) driving the fan, may includeany other suitable number or arrangement of shafts, spools, compressors,turbines, etc.

As discussed above, the turbofan engine 10 of the present disclosurealso provides pre-swirling flow forward a tip of the fan blade 40.Referring now also to FIG. 2 , a close-up, cross-sectional view of thefan section 14 and forward end of the turbomachine 16 of the exemplaryturbofan engine 10 of FIG. 1 is provided. In exemplary embodiments, theturbofan engine 10 includes an inlet pre-swirl vane located upstream ofthe plurality of fan blades 40 of the fan 38 and coupled to the outernacelle 50. For example, the inlet pre-swirl vane can be directlyattached to, indirectly attached to, or integrated into the outernacelle 50. More specifically, for the embodiment of FIGS. 1 and 2 , theinlet pre-swirl vane is configured as a plurality of part span inletguide vanes 100. The plurality of part span inlet guide vanes 100 areeach cantilevered from the outer nacelle 50 (such as from the inner wall52 of the outer nacelle 50) at a location forward of the plurality offan blades 40 of the fan 38 along the axial direction A and aft of theinlet 60 of the outer nacelle 50. More specifically, each of theplurality of part span inlet guide vanes 100 define an outer end 102along the radial direction R, and are coupled to the outer nacelle 50 atthe radially outer end 102 through a suitable connection means (notshown). For example, each of the plurality of part span inlet guidevanes 100 may be bolted to the inner wall 52 of the outer nacelle 50 atthe outer end 102, welded to the inner wall 52 of the outer nacelle 50at the outer end 102, or coupled to the outer nacelle 50 in any othersuitable manner at the outer end 102.

Referring still to FIG. 2 , in an exemplary embodiment, a nacelleassembly 80 of the present disclosure includes the outer nacelle 50 andthe inlet pre-swirl vane, e.g., a plurality of part span inlet guidevanes 100. Further, for the embodiment depicted, the plurality of partspan inlet guide vanes 100 extend generally along the radial direction Rfrom the outer end 102 to an inner end 104 (i.e., an inner end 104 alongthe radial direction R). Moreover, as will be appreciated, for theembodiment depicted, each of the plurality of part span inlet guidevanes 100 are unconnected with an adjacent part span inlet guide vane100 at the respective inner ends 104 (i.e., adjacent part span inletguide vanes 100 do not contact one another at the radially inner ends104, and do not include any intermediate connection members at theradially inner ends 104, such as a connection ring, strut, etc.). Morespecifically, for the embodiment depicted, each part span inlet guidevane 100 is completely supported by a connection to the outer nacelle 50at the respective outer end 102 (and not through any structureextending, e.g., between adjacent part span inlet guide vanes 100 at alocation inward of the outer end 102 along the radial direction R). Aswill be discussed below, such may reduce an amount of turbulencegenerated by the part span inlet guide vanes 100.

Moreover, as depicted, each of the plurality of part span inlet guidevanes 100 do not extend completely between the outer nacelle 50 and,e.g., the hub 48 of the turbofan engine 10. More specifically, for theembodiment depicted, each of the plurality of inlet guide vanes definean inlet guide vane (“IGV”) span 106 along the radial direction R, whichrefers to a measure along the radial direction R between the outer end102 and the inner end 104 of the part span inlet guide vane 100 at theleading edge 108 of the part span inlet guide vane 100. Each of theplurality of part span inlet guide vanes 100 further define a leadingedge 108 and a trailing edge 110. Similarly, it will be appreciated,that the plurality of fan blades 40 of the fan 38 define a fan bladespan 112 along the radial direction R, which refers to a measure alongthe radial direction R between a radially outer tip and a base of thefan blade 40 at the leading edge 114 of the respective fan blade 40.Each of the plurality of fan blades 40 of the fan 38 also defines aleading edge 114 and a trailing edge 116.

For the embodiment depicted, the IGV span 106 is at least about fivepercent of the fan blade span 112 and up to about fifty-five percent ofthe fan blade span 112. For example, in certain exemplary embodiments,the IGV span 106 may be between about fifteen percent of the fan bladespan 112 and about forty-five percent of the fan blade span 112, such asbetween about thirty percent of the fan blade span 112 and about fortypercent of the fan blade span 112.

Reference will now also be made to FIG. 3 , providing an axial view ofthe inlet 60 to the turbofan engine 10 of FIGS. 1 and 2 . As will beappreciated, the plurality of part span inlet guide vanes 100 of theturbofan engine 10 includes a relatively large number of part span inletguide vanes 100. For example, for the embodiment depicted, the pluralityof part span inlet guide vanes 100 of the turbofan engine 10 includesthirty-two part span inlet guide vanes 100. In other exemplaryembodiments, it is contemplated that the plurality of part span inletguide vanes 100 includes between about ten part span inlet guide vanes100 and about fifty part span inlet guide vanes 100. In furtherexemplary embodiments, it is contemplated that the plurality of partspan inlet guide vanes 100 includes between about twenty part span inletguide vanes 100 and about forty-five part span inlet guide vanes 100.Additionally, for the embodiment depicted, each of the plurality of partspan inlet guide vanes 100 are spaced substantially evenly (e.g.,equidistant) along the circumferential direction C. More specifically,each of the plurality of part span inlet guide vanes 100 defines acircumferential spacing 118 with an adjacent part span inlet guide vane100, with the circumferential spacing 118 being substantially equalbetween each adjacent part span inlet guide vane 100.

Although not depicted, in certain exemplary embodiments, the number ofpart span inlet guide vanes 100 may be substantially equal to the numberof fan blades 40 (FIG. 1 ) of the fan 38 (FIG. 1 ) of the turbofanengine 10. In other embodiments, however, the number of part span inletguide vanes 100 may be greater than the number of fan blades 40 of thefan 38 of the turbofan engine 10, or alternatively, may be less than thenumber of fan blades 40 of the fan 38 of the turbofan engine 10.

Further, it should be appreciated, that in other exemplary embodiments,the turbofan engine 10 may include any other suitable number of partspan inlet guide vanes 100 and/or circumferential spacing 118 of thepart span inlet guide vanes 100. For example, referring now briefly toFIG. 4 , an axial view of an inlet 60 to a turbofan engine 10 inaccordance with another exemplary embodiment of the present disclosureis provided. For the embodiment of FIG. 4 , the turbofan engine 10includes less than twenty part span inlet guide vanes 100. Morespecifically, for the embodiment of FIG. 4 , the turbofan engine 10includes at least eight part span inlet guide vanes 100, or morespecifically includes exactly eight part span inlet guide vanes 100.Additionally, for the embodiment of FIG. 4 , the plurality of part spaninlet guide vanes 100 are not substantially evenly spaced along thecircumferential direction C. For example, at least certain of theplurality of part span inlet guide vanes 100 define a firstcircumferential spacing 118A, while other of the plurality of part spaninlet guide vanes 100 define a second circumferential spacing 118B. Forthe embodiment depicted, the first circumferential spacing 118A is atleast about twenty percent greater than the second circumferentialspacing 118B, such as at least about twenty-five percent greater such asat least about thirty percent greater, such as up to about two hundredpercent greater. Notably, the circumferential spacing 118 refers to amean circumferential spacing between adjacent part span inlet guidevanes 100. The non-uniform circumferential spacing may, e.g., offsetstructure upstream of the part span inlet guide vanes 100.

Referring back to FIG. 3 , the outer nacelle 50 defines a radius NR. Forthe embodiment depicted, the plurality of part span inlet guide vanes100 are angled (e.g., tilted) at a first angle A1 with respect to theradius NR of the outer nacelle 50.

In this manner, the plurality of part span inlet guide vanes 100 areconfigured to direct incoming objects towards an outer radial portion ofthe turbofan engine 10 (FIG. 1 ). The plurality of part span inlet guidevanes 100 are configured to direct incoming objects away from the coreair flowpath 37 (FIG. 1 ) of the turbofan engine 10 and towards a bypassairflow passage 56 (FIG. 1 ). This provides a deflection mechanism thatfacilitates ingestion of an object into an outer radial portion of theturbofan engine 10 (FIG. 1 ) and thereby minimizing the chance that theobject travels to the core of the turbofan engine 10.

Furthermore, in this manner, the plurality of part span inlet guidevanes 100 are also configured to pre-swirl an airflow 58 (FIG. 1 )provided through the inlet 60 of the outer nacelle 50, upstream of theplurality of fan blades 40 (FIG. 1 ) of the fan 38 (FIG. 1 ). Asdiscussed herein, pre-swirling the airflow 58 provided through the inlet60 of the nacelle 50 prior to such airflow 58 reaching the plurality offan blades 40 of the fan 38 may reduce separation losses and/or shocklosses, allowing the fan 38 to operate with the relatively high fan tipspeeds described above with less losses in efficiency. Furthermore, inthis manner, the plurality of part span inlet guide vanes 100 are alsoconfigured to minimize flutter and maintain laminar air flow over thepart span inlet guide vanes 100 and minimize turbulence in the inlet 60.

In certain exemplary embodiments, the first angle A1 is betweenapproximately 2 degrees and approximately 45 degrees. In other exemplaryembodiments, it is contemplated that the first angle A1 is between otherranges for a particular application.

In the exemplary embodiment depicted in FIG. 3 , the plurality of partspan inlet guide vanes 100 are angled at the first angle A1 with respectto the radius NR of the outer nacelle 50 in a clockwise direction CW,i.e., the same direction that the fan blades 40 (FIG. 2 ) rotate in.

In an exemplary embodiment, the plurality of part span inlet guide vanes100 are each angled at the same first angle A1 with respect to theradius NR of the outer nacelle 50. In other exemplary embodiments, theplurality of part span inlet guide vanes 100 are angled at differentfirst angles A1 with respect to the radius NR of the outer nacelle 50 aswill be described in more detail below.

Referring now to FIG. 5 , an axial view of the inlet 60 to the turbofanengine 10 of FIGS. 1 and 2 , according to another exemplary embodiment,is provided. Referring to FIG. 5 , in another exemplary embodiment, theplurality of part span inlet guide vanes 100 are angled at the firstangle A1 with respect to the radius NR of the outer nacelle 50 in acounterclockwise direction CCW, i.e., the opposite direction that thefan blades 40 (FIG. 2 ) rotate in.

Referring now back to FIG. 2 , as described above, each of the pluralityof part span inlet guide vanes 100 are configured to pre-swirl anairflow 58 provided through the inlet 60 of the nacelle 50, upstream ofthe plurality of fan blades 40 of the fan 38. As discussed above,pre-swirling the airflow 58 provided through the inlet 60 of the nacelle50 prior to such airflow 58 reaching the plurality of fan blades 40 ofthe fan 38 may reduce separation losses and/or shock losses, allowingthe fan 38 to operate with the relatively high fan tip speeds describedabove with less losses of in efficiency.

For example, referring first to FIG. 6 , a cross-sectional view of onepart span inlet guide vane 100 along the span of the part span inletguide vanes 100, as indicated by Line 6-6 in FIG. 2 , is provided. As isdepicted, the part span inlet guide vane 100 is configured generally asan airfoil having a pressure side 120 and an opposite suction side 122,and extending between the leading edge 108 and the trailing edge 110along a camber line 124. Additionally, the part span inlet guide vane100 defines a chord line 126 extending directly from the leading edge108 to the trailing edge 110. The chord line 126 of the part span inletguide vane 100 is angled (e.g., twisted) at a second angle or angle ofattack 128 with respect to the longitudinal axis 12 of the outer nacelle50 (FIG. 2 ). For example, the chord line 126 defines a second angle orangle of attack 128 with an airflow direction 129 of the airflow 58through the inlet 60 (FIG. 2 ) of the nacelle 50. Notably, for theembodiment depicted, the airflow direction 129 is substantially parallelto the axial direction A and the longitudinal axis 12 of the outernacelle 50 of the turbofan engine 10. For the embodiment depicted, theangle of attack 128 at the location depicted along the span 106 of thepart span inlet guide vanes 100 is at least approximately five degreesand up to approximately thirty-five degrees. For example, in certainembodiments, the angle of attack 128 at the location depicted along thespan 106 of the part span inlet guide vane 100 may be between about tendegrees and about thirty degrees, such as between about fifteen degreesand about twenty-five degrees.

Additionally, the part span inlet guide vane 100, at the locationdepicted along the span 106 (FIG. 2 ) of the part span inlet guide vane100 defines a local swirl angle 130 at the trailing edge 110. The “swirlangle” at the trailing edge 110 of the part span inlet guide vane 100,as used herein, refers to an angle between the airflow direction 129 ofthe airflow 58 through the inlet 60 (FIG. 2 ) of the nacelle 50 (FIG. 2) and a reference line 132 defined by a trailing edge section of thepressure side 120 of the part span inlet guide vane 100. Morespecifically, the reference line 132 is defined by the aft twentypercent of the pressure side 120, as measured along the chord line 126.Notably, when the aft twenty percent the pressure side 120 defines acurve, the reference line 132 may be a straight-line average fit of suchcurve (e.g., using least mean squares).

Further, it will be appreciated, that a maximum swirl angle 130 refersto the highest swirl angle 130 along the span 106 (FIG. 2 ) of the partspan inlet guide vane 100. For the embodiment depicted, the maximumswirl angle 130 is defined proximate the radially outer end 102 (FIG. 2) of the part span inlet guide vane 100 (e.g., at the outer ten percentof the span 106 of the part span inlet guide vanes 100), as isrepresented by the cross-section depicted in FIG. 6 . For the embodimentdepicted, the maximum swirl angle 130 of each part span inlet guide vane100 at the trailing edge 110 is between approximately five degrees andapproximately thirty-five degrees. For example, in certain exemplaryembodiments, the maximum swirl angle 130 of each part span inlet guidevane 100 at the trailing edge 110 may be between twelve degrees andtwenty-five degrees.

Moreover, it should be appreciated that for the embodiment of FIG. 2 ,the local swirl angle 130 increases from the radially inner end 104(FIG. 2 ) to the radially outer end 102 (FIG. 2 ) of each part spaninlet guide vane 100. For example, referring now also to FIG. 7 , across-sectional view of a part span inlet guide vane 100 at a locationradially inward from the cross-section viewed in FIG. 6 , as indicatedby Line 7-7 in FIG. 2 , is provided. As is depicted in FIG. 7 , and asstated above, the part span inlet guide vane 100 defines the pressureside 120, the suction side 122, the leading edge 108, the trailing edge110, the camber line 124, and chord line 126. Further, the second angleor angle of attack 128 defined by the chord line 126 and the airflowdirection 129 of the airflow 58 through the inlet 60 of the nacelle 50at the location along the span 106 depicted in FIG. 7 is less than theangle of attack 128 at the location along the span 106 depicted in FIG.6 (e.g., may be at least about twenty percent less, such as at leastabout fifty percent less, such as up to about one hundred percent less).Additionally, the part span inlet guide vane 100 defines a local swirlangle 130 at the trailing edge 110 at the location along the span 106 ofthe part span inlet guide vane 100 proximate the inner end 104, asdepicted in FIG. 7 . As stated above, the local swirl angle 130increases from the radially inner end 104 to the radially outer end 102of each part span inlet guide vanes 100. Accordingly, the local swirlangle 130 proximate the outer end 102 (see FIG. 6 ) is greater than thelocal swirl angle 130 proximate the radially inner end 104 (see FIG. 7 ;e.g., the radially inner ten percent of the span 106). For example, thelocal swirl angle 130 may approach zero degrees (e.g., may be less thanabout five degrees, such as less than about two degrees) at the radiallyinner end 104.

Notably, including part span inlet guide vanes 100 of such aconfiguration may reduce an amount of turbulence at the radially innerend 104 (FIG. 2 ) of each respective part span inlet guide vane 100.Additionally, such a configuration may provide a desired amount ofpre-swirl at the radially outer ends of the plurality of fan blades 40(FIG. 2 ) of the fan 38 (FIG. 2 ) (where the speed of the fan blades 40is the greatest) to provide a desired reduction in flow separationand/or shock losses that may otherwise occur due to a relatively highspeed of the plurality of fan blades 40 at the fan tips during operationof the turbofan engine 10 (FIG. 2 ).

Referring now to FIG. 8 , an axial view of the inlet 60 to the turbofanengine 10 of FIGS. 1 and 2 , according to another exemplary embodiment,is provided. Referring to FIG. 8 , in another exemplary embodiment, theplurality of part span inlet guide vanes 100 are angled at differentfirst angles A1 with respect to the radius NR of the outer nacelle 50.

The outer nacelle 50 includes a top portion 210, a bottom portion 212, afirst side portion 214, and a second side portion 216. In an exemplaryembodiment, a first portion, e.g., the top portion 210, of the part spaninlet guide vanes 100 are angled at a first angle A1 with respect to theradius NR of the outer nacelle 50. For example, the part span inletguide vanes 100 are angled at a first angle A1 of five degrees or sevendegrees with respect to the radius NR of the outer nacelle 50 at the topportion 210. Furthermore, a second portion, e.g., the first side portion214, of the part span inlet guide vanes 100 are angled at a differentangle with respect to the radius NR of the outer nacelle 50. Forexample, the part span inlet guide vanes 100 are angled at a first angleA1 of ten degrees or fifteen degrees with respect to the radius NR ofthe outer nacelle 50 at the first side portion 214. In such exemplaryembodiments, the plurality of part span inlet guide vanes 100 are angledat different first angles A1 with respect to the radius NR of the outernacelle 50 along the circumferential direction C.

In such an exemplary embodiment, circumferential variation in the angleof tilt, e.g., the first angle A1 with respect to the radius NR of theouter nacelle 50, can address many issues including cross winds, highangle of attack maneuvers such as takeoff, and engine installation onone side of an aircraft or another. For example, cross winds are morelikely to affect the 3 or 9 o'clock positions on the engine, e.g., thesecond side portion 216 and the first side portion 214, which may make adifferent angle of tilt more desirable at the 3/9 o'clock positions thanat the 6/12 o'clock positions on the engine, e.g., the bottom portion212 and the top portion 210. In addition, installation on one side ofthe aircraft or the other may make cross wind effects more pronounced onthe side of the engine that is further from the fuselage. Likewise, ahigh angle of attack may make differing tilt angles at the 6/12 o'clockpositions desirable, e.g., the bottom portion 212 and the top portion210. For these reasons, it is contemplated that the first angles A1 withrespect to the radius NR of the outer nacelle 50 may vary for aparticular application and may be different at the top portion 210, thebottom portion 212, the first side portion 214, and/or the second sideportion 216 of the outer nacelle 50.

Referring now to FIG. 9 , an axial view of the inlet 60 to the turbofanengine 10 of FIGS. 1 and 2 , according to another exemplary embodiment,is provided. Referring to FIG. 9 , in another exemplary embodiment, theplurality of part span inlet guide vanes 100 are angled at differentfirst angles A1 with respect to the radius NR of the outer nacelle 50.

The outer nacelle 50 includes a top portion 210, a bottom portion 212, afirst side portion 214, and a second side portion 216. In an exemplaryembodiment, the part span inlet guide vanes 100 are angled at differentfirst angles A1 with respect to the radius NR of the outer nacelle 50 atthe top portion 210 and the bottom portion 212. For example, the topportion 210 of the part span inlet guide vanes 100 are angled at a firstangle A1 with respect to the radius NR of the outer nacelle 50. Forexample, the part span inlet guide vanes 100 are angled at a first angleA1 of five degrees or seven degrees with respect to the radius NR of theouter nacelle 50 at the top portion 210. Furthermore, the bottom portion212 of the part span inlet guide vanes 100 are angled at a differentangle with respect to the radius NR of the outer nacelle 50. Forexample, the part span inlet guide vanes 100 are angled at a first angleA1 of ten degrees with respect to the radius NR of the outer nacelle 50at the bottom portion 212. In such exemplary embodiments, the pluralityof part span inlet guide vanes 100 are angled at different first anglesA1 with respect to the radius NR of the outer nacelle 50 at the topportion 210 and the bottom portion 212.

In such an exemplary embodiment, circumferential variation in the angleof tilt, e.g., the first angle A1 with respect to the radius NR of theouter nacelle 50, can address many issues including cross winds, highangle of attack maneuvers such as takeoff, and engine installation onone side of an aircraft or another. For example, cross winds are morelikely to affect the 3 or 9 o'clock positions on the engine, e.g., thesecond side portion 216 and the first side portion 214, which may make adifferent angle of tilt more desirable at the 3/9 o'clock positions thanat the 6/12 o'clock positions on the engine, e.g., the bottom portion212 and the top portion 210. In addition, installation on one side ofthe aircraft or the other may make cross wind effects more pronounced onthe side of the engine that is further from the fuselage. Likewise, ahigh angle of attack may make differing tilt angles at the 6/12 o'clockpositions desirable, e.g., the bottom portion 212 and the top portion210. For these reasons, it is contemplated that the first angles A1 withrespect to the radius NR of the outer nacelle 50 may vary for aparticular application and may be different at the top portion 210, thebottom portion 212, the first side portion 214, and/or the second sideportion 216 of the outer nacelle 50.

Referring now to FIG. 10 , an axial view of the inlet 60 to the turbofanengine 10 of FIGS. 1 and 2 , according to another exemplary embodiment,is provided. Referring to FIG. 10 , in another exemplary embodiment, oneof the part span inlet guide vanes 100 are angled at different angleswith respect to the radius NR of the outer nacelle 50 at differentportions of the part span inlet guide vane 100.

For example, the part span inlet guide vane 100 is angled at a firstangle A1 with respect to the radius NR of the outer nacelle 50 a firstlocation 250 of the part span inlet guide vane 100 and the part spaninlet guide vane 100 is angled at a second angle A2 with respect to theradius NR of the outer nacelle 50 at a second location 252 of the partspan inlet guide vane 100. In such an embodiment, the first angle A1 isdifferent than the second angle A2. For example, in the embodimentdepicted in FIG. 10 , the first angle A1 is greater than the secondangle A2.

In such an exemplary embodiment, varying the angle of tilt, e.g., thefirst angle A1 and the second angle A2, from the base of the part spaninlet guide vane 100, e.g., the outer end 102 (FIG. 2 ), to the tip ofthe part span inlet guide vane 100, e.g., the inner end 104 (FIG. 2 ),can address varying levels of turbulence due to boundary layer effects,as you get closer to/further from the wall 52 (FIG. 2 ) of the outernacelle 50. The angle of tilt and sweep may also be used to improve thematching of the swirl that is imparted to the incoming air with thevarying linear speed of the fan blade 40 (FIG. 2 ) as you get closerto/further from the axis 12 (FIG. 1 ) of the engine 10. The linear speedof the fan blade 40 (FIG. 2 ) is greatest at the tip, and slowest at theroot of the fan blade 40 (FIG. 2 ), even though the rotational velocityis equal.

Referring now to FIG. 11 , an axial view of the inlet 60 to the turbofanengine 10 of FIGS. 1 and 2 , according to another exemplary embodiment,is provided. Referring to FIG. 11 , in another exemplary embodiment, oneof the part span inlet guide vanes 100 are angled at different angleswith respect to the radius NR of the outer nacelle 50 at differentlocations of the part span inlet guide vane 100.

For example, the part span inlet guide vane 100 is angled at a firstangle A1 with respect to the radius NR of the outer nacelle 50 a firstlocation 250 of the part span inlet guide vane 100 and the part spaninlet guide vane 100 is angled at a second angle A2 with respect to theradius NR of the outer nacelle 50 at a second location 252 of the partspan inlet guide vane 100. In such an embodiment, the first angle A1 isdifferent than the second angle A2. For example, in the embodimentdepicted in FIG. 11 , the first angle A1 is less than the second angleA2.

In such an exemplary embodiment, varying the angle of tilt, e.g., thefirst angle A1 and the second angle A2, from the base of the part spaninlet guide vane 100, e.g., the outer end 102 (FIG. 2 ), to the tip ofthe part span inlet guide vane 100, e.g., the inner end 104 (FIG. 2 ),can address varying levels of turbulence due to boundary layer effects,as you get closer to/further from the wall 52 (FIG. 2 ) of the outernacelle 50. The angle of tilt and sweep may also be used to improve thematching of the swirl that is imparted to the incoming air with thevarying linear speed of the fan blade 40 (FIG. 2 ) as you get closerto/further from the axis 12 (FIG. 1 ) of the engine 10. The linear speedof the fan blade 40 (FIG. 2 ) is greatest at the tip, and slowest at theroot of the fan blade 40 (FIG. 2 ), even though the rotational velocityis equal.

Further aspects of the disclosure are provided by the subject matter ofthe following clauses:

A turbofan engine comprising: a fan comprising a plurality of fanblades; a turbomachine operably coupled to the fan for driving the fan,the turbomachine comprising a compressor section, a combustion section,and a turbine section in serial flow order and together defining a coreair flowpath; a nacelle surrounding and at least partially enclosing thefan, the nacelle defining a radius and a longitudinal axis; and an inletpre-swirl vane located upstream of the plurality of fan blades anddefining a chord, the inlet pre-swirl vane coupled to the nacelle,wherein the inlet pre-swirl vane is angled at a first angle with respectto the radius of the nacelle, and wherein the chord of the inletpre-swirl vane is angled at a second angle with respect to thelongitudinal axis of the nacelle.

The turbofan engine of any preceding clause, wherein the inlet pre-swirlvane is angled at the first angle with respect to the radius of thenacelle in a clockwise direction from an inlet of the nacelle.

The turbofan engine of any preceding clause, wherein the inlet pre-swirlvane is angled at the first angle with respect to the radius of thenacelle in a counterclockwise direction from an inlet of the nacelle.

The turbofan engine of any preceding clause, wherein the first angle isbetween approximately 2 degrees and approximately 45 degrees.

The turbofan engine of any preceding clause, wherein the second angle isbetween approximately 5 degrees and approximately 35 degrees.

The turbofan engine of any preceding clause, wherein the inlet pre-swirlvane is one of a plurality of part span inlet guide vanes extending fromthe nacelle upstream of the plurality of fan blades and aft of an inletof the nacelle.

The turbofan engine of any preceding clause, wherein each of theplurality of part span inlet guide vanes is angled at the same firstangle with respect to the radius of the nacelle.

The turbofan engine of any preceding clause, wherein a first portion ofeach of the plurality of part span inlet guide vanes is angled at thefirst angle with respect to the radius of the nacelle, wherein a secondportion of each of the plurality of part span inlet guide vanes isangled at a third angle with respect to the radius of the nacelle, andwherein the first angle is different than the third angle.

The turbofan engine of any preceding clause, wherein the nacelleincludes a top portion, a bottom portion, a first side portion, and asecond side portion, wherein a first portion of each of the plurality ofpart span inlet guide vanes is angled at the first angle with respect tothe radius of the nacelle at the top portion, wherein a second portionof each of the plurality of part span inlet guide vanes is angled at athird angle with respect to the radius of the nacelle at the bottomportion, and wherein the first angle is different than the third angle.

A nacelle assembly for a turbofan engine, the turbofan engine comprisinga fan including a plurality of fan blades, the nacelle assemblycomprising: a nacelle surrounding and at least partially enclosing thefan, the nacelle defining a radius and a longitudinal axis; and an inletpre-swirl vane located upstream of the plurality of fan blades anddefining a chord, the inlet pre-swirl vane coupled to the nacelle,wherein the inlet pre-swirl vane is angled at a first angle with respectto the radius of the nacelle, and wherein the chord of the inletpre-swirl vane is angled at a second angle with respect to thelongitudinal axis of the nacelle.

The nacelle assembly of any preceding clause, wherein the inletpre-swirl vane is angled at the first angle with respect to the radiusof the nacelle in a clockwise direction from an inlet of the nacelle.

The nacelle assembly of any preceding clause, wherein the inletpre-swirl vane is angled at the first angle with respect to the radiusof the nacelle in a counterclockwise direction from an inlet of thenacelle.

The nacelle assembly of any preceding clause, wherein the first angle isbetween approximately 2 degrees and approximately 45 degrees.

The nacelle assembly of any preceding clause, wherein the second angleis between approximately 5 degrees and approximately 35 degrees.

The nacelle assembly of any preceding clause, wherein the inletpre-swirl vane is one of a plurality of part span inlet guide vanesextending from the nacelle upstream of the plurality of fan blades andaft of an inlet of the nacelle.

The nacelle assembly of any preceding clause, wherein each of theplurality of part span inlet guide vanes is angled at the same firstangle with respect to the radius of the nacelle.

The nacelle assembly of any preceding clause, wherein a first portion ofeach of the plurality of part span inlet guide vanes is angled at thefirst angle with respect to the radius of the nacelle, wherein a secondportion of each of the plurality of part span inlet guide vanes isangled at a third angle with respect to the radius of the nacelle, andwherein the first angle is different than the third angle.

The nacelle assembly of any preceding clause, wherein the nacelleincludes a top portion, a bottom portion, a first side portion, and asecond side portion, wherein a first portion of each of the plurality ofpart span inlet guide vanes is angled at the first angle with respect tothe radius of the nacelle at the top portion, wherein a second portionof each of the plurality of part span inlet guide vanes is angled at athird angle with respect to the radius of the nacelle at the bottomportion, and wherein the first angle is different than the third angle.

A nacelle assembly for a turbofan engine, the turbofan engine comprisinga fan including a plurality of fan blades, the nacelle assemblycomprising: a nacelle surrounding and at least partially enclosing thefan, the nacelle defining a radius; and an inlet pre-swirl vane locatedupstream of the plurality of fan blades, the inlet pre-swirl vanecoupled to the nacelle, wherein the inlet pre-swirl vane is angled at afirst angle with respect to the radius of the nacelle at a firstlocation of the inlet pre-swirl vane, wherein the inlet pre-swirl vaneis angled at a second angle with respect to the radius of the nacelle ata second location of the inlet pre-swirl vane, and wherein the firstangle is different than the second angle.

The nacelle assembly of any preceding clause, wherein the nacelledefines a longitudinal axis, wherein the inlet pre-swirl vane defines achord, and wherein the chord of the inlet pre-swirl vane is angled at athird angle with respect to the longitudinal axis of the nacelle.

This written description uses examples to disclose the disclosure,including the best mode, and also to enable any person skilled in theart to practice the disclosure, including making and using any devicesor systems and performing any incorporated methods. The patentable scopeof the disclosure is defined by the claims, and may include otherexamples that occur to those skilled in the art. Such other examples areintended to be within the scope of the claims if they include structuralelements that do not differ from the literal language of the claims, orif they include equivalent structural elements with insubstantialdifferences from the literal languages of the claims.

While this disclosure has been described as having exemplary designs,the present disclosure can be further modified within the scope of thisdisclosure. This application is therefore intended to cover anyvariations, uses, or adaptations of the disclosure using its generalprinciples. Further, this application is intended to cover suchdepartures from the present disclosure as come within known or customarypractice in the art to which this disclosure pertains and which fallwithin the limits of the appended claims.

1. A turbofan engine comprising: a fan comprising a plurality of fanblades; a turbomachine operably coupled to the fan for driving the fan,the turbomachine comprising a compressor section, a combustion section,and a turbine section in serial flow order and together defining a coreair flowpath; a nacelle surrounding and at least partially enclosing thefan, the nacelle defining a radius and a longitudinal axis; and an inletpre-swirl vane located upstream of the plurality of fan blades anddefining a chord, the inlet pre-swirl vane coupled to the nacelle,wherein the inlet pre-swirl vane is angled at a first angle with respectto the radius of the nacelle, and wherein the chord of the inletpre-swirl vane is angled at a second angle with respect to thelongitudinal axis of the nacelle.
 2. The turbofan engine of claim 1,wherein the inlet pre-swirl vane is angled at the first angle withrespect to the radius of the nacelle in a clockwise direction from aninlet of the nacelle.
 3. The turbofan engine of claim 1, wherein theinlet pre-swirl vane is angled at the first angle with respect to theradius of the nacelle in a counterclockwise direction from an inlet ofthe nacelle.
 4. The turbofan engine of claim 1, wherein the first angleis between approximately 2 degrees and approximately 45 degrees.
 5. Theturbofan engine of claim 1, wherein the second angle is betweenapproximately 5 degrees and approximately 35 degrees.
 6. The turbofanengine of claim 1, wherein the inlet pre-swirl vane is one of aplurality of part span inlet guide vanes extending from the nacelleupstream of the plurality of fan blades and aft of an inlet of thenacelle.
 7. The turbofan engine of claim 6, wherein each of theplurality of part span inlet guide vanes is angled at the same firstangle with respect to the radius of the nacelle.
 8. The turbofan engineof claim 6, wherein a first portion of each of the plurality of partspan inlet guide vanes is angled at the first angle with respect to theradius of the nacelle, wherein a second portion of each of the pluralityof part span inlet guide vanes is angled at a third angle with respectto the radius of the nacelle, and wherein the first angle is differentthan the third angle.
 9. The turbofan engine of claim 6, wherein thenacelle includes a top portion, a bottom portion, a first side portion,and a second side portion, wherein a first portion of each of theplurality of part span inlet guide vanes is angled at the first anglewith respect to the radius of the nacelle at the top portion, wherein asecond portion of each of the plurality of part span inlet guide vanesis angled at a third angle with respect to the radius of the nacelle atthe bottom portion, and wherein the first angle is different than thethird angle.
 10. A nacelle assembly for a turbofan engine, the turbofanengine comprising a fan including a plurality of fan blades, the nacelleassembly comprising: a nacelle surrounding and at least partiallyenclosing the fan, the nacelle defining a radius and a longitudinalaxis; and an inlet pre-swirl vane located upstream of the plurality offan blades and defining a chord, the inlet pre-swirl vane coupled to thenacelle, wherein the inlet pre-swirl vane is angled at a first anglewith respect to the radius of the nacelle, and wherein the chord of theinlet pre-swirl vane is angled at a second angle with respect to thelongitudinal axis of the nacelle.
 11. The nacelle assembly of claim 10,wherein the inlet pre-swirl vane is angled at the first angle withrespect to the radius of the nacelle in a clockwise direction from aninlet of the nacelle.
 12. The nacelle assembly of claim 10, wherein theinlet pre-swirl vane is angled at the first angle with respect to theradius of the nacelle in a counterclockwise direction from an inlet ofthe nacelle.
 13. The nacelle assembly of claim 10, wherein the firstangle is between approximately 2 degrees and approximately 45 degrees.14. The nacelle assembly of claim 10, wherein the second angle isbetween approximately 5 degrees and approximately 35 degrees.
 15. Thenacelle assembly of claim 10, wherein the inlet pre-swirl vane is one ofa plurality of part span inlet guide vanes extending from the nacelleupstream of the plurality of fan blades and aft of an inlet of thenacelle.
 16. The nacelle assembly of claim 15, wherein each of theplurality of part span inlet guide vanes is angled at the same firstangle with respect to the radius of the nacelle.
 17. The nacelleassembly of claim 15, wherein a first portion of each of the pluralityof part span inlet guide vanes is angled at the first angle with respectto the radius of the nacelle, wherein a second portion of each of theplurality of part span inlet guide vanes is angled at a third angle withrespect to the radius of the nacelle, and wherein the first angle isdifferent than the third angle.
 18. The nacelle assembly of claim 15,wherein the nacelle includes a top portion, a bottom portion, a firstside portion, and a second side portion, wherein a first portion of eachof the plurality of part span inlet guide vanes is angled at the firstangle with respect to the radius of the nacelle at the top portion,wherein a second portion of each of the plurality of part span inletguide vanes is angled at a third angle with respect to the radius of thenacelle at the bottom portion, and wherein the first angle is differentthan the third angle.
 19. (canceled)
 20. (canceled)
 21. The turbofanengine of claim 1, wherein the inlet pre-swirl vane includes an outerend and an inner end, wherein the inner end is radially inward from theouter end, and wherein the second angle of the chord at the outer end isgreater than the second angle of the chord at the inner end.
 22. Theturbofan engine of claim 1, wherein the inlet pre-swirl vane includes anouter end and an inner end, wherein the inner end is radially inwardfrom the outer end, and wherein the second angle of the chord increasesfrom the inner end to the outer end.